JPH07217404A - Assembling method for turbine stationary blade - Google Patents

Abstract

PURPOSE:To assemble a turbine stationary blade without thermally influcencing on a nozzle. CONSTITUTION:A porous cooling plate 23 is soldered so as to cover a cooling groove 22 on an outer band 19 of a turbine stationary blade main body 21. The soldered portion is numbered 36. An insert 24 inserted into a nozzle hollow portion 32 is welded and fixed to a lip 35 formed at the cooling plate 23. The selded portion is numbered 37. The turbine stationary blade is thus assembled without influencing on a nozzle 18. The nozzle 18, if being made of monocrystal material, is prevented from being crushed in its monocrystallized condition.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a method of assembling a turbine vane.

[0002]

5 and 6 show an example of a gas turbine engine such as a jet engine.

A low pressure compressor drive shaft 2 is rotatably disposed at the axial center of a cylindrical engine body 1, a low pressure compressor 3 is attached to the tip of the low pressure compressor drive shaft 2, and a low pressure compressor drive is performed. A low pressure turbine 4 is attached to the rear end of the shaft 2.

A hollow high-pressure compressor drive shaft 5 is rotatably fitted on the outer periphery of the intermediate portion of the low-pressure compressor drive shaft 2, a high-pressure compressor 6 is attached to the tip of the high-pressure compressor drive shaft 5, and high-pressure compression is performed. A high pressure turbine 7 is attached to the rear end of the machine drive shaft 5, and a combustor 8 is provided inside the engine body 1 at an intermediate position between the high pressure compressor 6 and the high pressure turbine 7.

Further, a low-pressure compressor blade 10 projecting outward from the engine body 1 is attached to a disk 9 at the tip of the low-pressure compressor 3, and an outer cylinder 11 surrounding the low-pressure compressor blade 10 is attached to the outer periphery of the engine body 1. The fan outlet guide vanes 12 are provided between the engine body 1 and the outer cylinder 11.

Reference numeral 13 is air, 14 is combustion gas generated in the combustor 8, 15 is a bypass passage formed between the engine body 1 and the outer cylinder 11, and 16 is a high pressure turbine rotor blade of the high pressure turbine 7. , 17 are high-pressure turbine stationary blades.

Then, the fuel is supplied to the combustor 8 provided inside the engine body 1, and the air 13 is mixed with the fuel and burned.

Then, the combustion gas 14 generated by the combustion in the combustor 8 flows rearward inside the engine body 1 to rotate the high pressure turbine 7 and the low pressure turbine 4 provided behind the combustor 8, and thereafter. The thrust is generated by being injected from the rear part of the engine body 1.

When the low pressure turbine 4 is rotated,
The low pressure compressor 3 is driven via the low pressure compressor drive shaft 2,
The low-pressure compressor moving blade 10 attached to the disk 9 of the low-pressure compressor 3 rotates to suck air 13 into the outer cylinder 11, and a part of the sucked air 13 is introduced into the engine body 1 to lower the pressure. It is compressed by the turbine 4.

The air 13 compressed by the low pressure turbine 4 is
It is compressed to a high pressure by the high-pressure compressor 6 driven via the high-pressure compressor drive shaft 5 by the rotation of the high-pressure turbine 7.

The air 13 compressed by the high pressure turbine 7 is
It enters the combustor 8 and is used for burning fuel as described above.

On the other hand, the low pressure compressor moving blade 10 causes the outer cylinder 1 to move.
The remaining air 13 introduced into the inside 1 flows through the bypass passage 15 between the engine body 1 and the outer cylinder 11, is rectified by the fan outlet guide vanes 12, and then injected from the rear portion of the outer cylinder 11 to generate thrust. It is supposed to be generated.

The high-pressure turbine stationary blade 17 placed at the outlet of the combustor 8 under the severest temperature conditions is shown in FIGS.
As shown in FIG. 2, the outer band 19 and the inner band 2 for attaching and fixing to the engine body 1 side and the high pressure compressor drive shaft 5 side are attached to both ends of the hollow porous nozzle 18 having a blade shape.
A turbine vane main body 21 having 0, a cooling groove 22 is formed around the opening 31 of the nozzle hollow portion 32 in the outer band 19 of the turbine vane main body 21, and the outer band 19 covers the cooling groove 22. The porous cooling plate 23 is fixed by welding and the nozzle hollow portion 3
A hollow porous insert 24 is inserted into the insert 2, and the end portion of the insert 24 and the opening 31 portion of the outer band 19 are fixed by welding.

In the drawing, 25 is a nozzle hole formed in the nozzle 18, 26 is a nozzle hole formed in the outer band 19, 27 is a nozzle hole formed in the cooling plate 23, and 28 is formed in the insert 24. Nozzle hole, 29
Is a welded portion between the outer band 19 and the outer edge portion 33 and the inner edge portion 34 of the cooling plate 23, and 30 is a welded portion between the opening 31 portion of the nozzle 18 and the insert 24.

For the high-pressure turbine stationary blade 17 having the above structure,
As shown in FIG. 6, the air 13 enters the inside of the insert 24 provided in the nozzle 18 from the outer band 19 side and the inner band 20 side, and blows out from the nozzle hole 28 formed in the insert 24 to clean the inner surface of the nozzle 18. Collision cooling.

Air 13 with the inner surface of the nozzle 18 impingingly cooled
Is then ejected from the nozzle holes 25 formed in the nozzle 18 to the outside of the nozzle 18, flows in a film shape along the outer periphery of the nozzle 18, and film-cools the nozzle 18.

On the outer band 19 side, a part of the air has a nozzle hole 27 formed in the cooling plate 23.
Enters the cooling groove portion 22 formed in the outer band 19 to cool the cooling groove portion 22 with the film, and is jetted toward the outside of the nozzle 18 from the nozzle hole 26 formed in the outer band 19 to form a film on the outer periphery of the nozzle 18. Cooling.

Although the nozzle 18, the insert 24, etc. are made of a material such as a nickel-base heat resistant alloy, the nozzle 18 which is directly exposed to the high temperature combustion gas 14 is used.
Is high temperature creep strength, that is, so that cracks do not form between the crystals that make up the material at high temperature,
In particular, a single crystallized nickel-base heat resistant alloy is used.

[0019]

However, since the cooling plate 23 and the insert 24 are directly welded and fixed to the nozzle 18 using the above-mentioned single crystal material, the nozzle 18 is heated by heat during welding. There was a problem that the material was changed under the influence of the influence and the single crystallinity was partially destroyed.

In view of the above situation, it is an object of the present invention to provide a method for assembling a turbine vane that can be assembled without exerting a thermal influence on the nozzle.

[0021]

DISCLOSURE OF THE INVENTION The present invention is directed to a turbine vane main body comprising an outer band and an inner band for mounting and fixing at both ends of a hollow porous nozzle having a blade shape, and a nozzle hollow in the outer band. The porous cooling plate is brazed so as to cover the cooling groove formed around the opening of the portion, and the inner edge portion of the cooling plate along the opening is bent outward to form the lip portion. The invention relates to a method for assembling a turbine vane, characterized in that a hollow porous insert inserted into the nozzle hollow portion is welded and fixed to the lip portion.

[0022]

The operation of the present invention is as follows.

A porous cooling plate is brazed so as to cover the cooling groove portion in the outer band of the turbine vane body, and an insert inserted into the hollow portion of the nozzle is welded and fixed to a lip portion formed on the cooling plate. The turbine vane is assembled without giving a thermal effect to the nozzle, and even if the nozzle is a single crystal material, the single crystallinity of the nozzle is prevented from being destroyed.

[0024]

Embodiments of the present invention will be described below with reference to the drawings.

1 to 4 show an embodiment of the present invention.

Since the basic structure of the jet engine itself is the same as that of FIGS.
Please refer to FIG.

As shown in FIGS. 7 to 10, the outer band 19 and the inner band for attaching and fixing to the engine body 1 side and the high-pressure compressor drive shaft 5 side are attached to both ends of the hollow porous nozzle 18 having a blade shape. A turbine vane body 21 having a band 20 is provided, and the outer band 1 of the turbine vane body 21 is provided.
9, the cooling groove 2 is formed around the opening 31 of the nozzle 18.
2, a porous cooling plate 23 that covers the cooling groove portion 22 is brazed and fixed to the outer band 19.

The insert 24 is inserted into the lip portion 35 formed by inserting the hollow porous insert 24 into the nozzle hollow portion 32 of the nozzle 18 and bending the inner edge portion 34 along the opening 31 of the cooling plate 23. The high pressure turbine stationary blade 17 is configured by welding and fixing.

In the figure, 25 is a nozzle hole formed in the nozzle 18, 26 is a nozzle hole formed in the outer band 19, 27 is a nozzle hole formed in the cooling plate 23, and 28 is formed in the insert 24. Nozzle hole, 36
Is a brazed portion between the outer band 19 and the cooling plate 23, and 37 is a welded portion 37 between the inner edge portion 34 of the cooling plate 23 and the insert 24.

The nozzle 18 and the insert 24 are made of a material such as a nickel-base heat-resistant alloy, and the nozzle 18 which comes into direct contact with the high temperature combustion gas 14 has a high temperature creep strength, that is, a high temperature. In particular, a single crystallized nickel-base heat-resistant alloy is used so that no cracks are formed between the crystals of the material below.

For the high-pressure turbine stationary blade 17 having the above structure,
The air 13 enters the inside of the insert 24 provided in the nozzle 18 from the outer band 19 side and the inner band 20 side, blows out from the nozzle hole 28 formed in the insert 24, and impinges and cools the inner surface of the nozzle 18.

The air 13 in which the inner surface of the nozzle 18 is collision-cooled
Is then ejected from the nozzle hole 25 formed in the nozzle 18 to the outside of the nozzle 18, and flows in a film shape along the outer periphery of the nozzle 18 to film-cool the nozzle 18.

On the outer band 19 side, part of the air is a nozzle hole 27 formed in the cooling plate 23.
Enters the cooling groove portion 22 formed in the outer band 19 to cool the cooling groove portion 22 with the film, and is jetted toward the outside of the nozzle 18 from the nozzle hole 26 formed in the outer band 19 to form a film on the outer periphery of the nozzle 18. Cooling.

Here, the nozzle 18 that comes into direct contact with the hot combustion gas 14 has a high high temperature creep strength, that is, does not crack between the crystals of the material at a high temperature. Although a single crystallized nickel-base heat-resistant alloy is used, in the present invention, the cooling plate 23 is fixed to the single crystal nozzle 18 by brazing with less heat effect, and the lip portion formed on the cooling plate 23 is fixed. Since the insert 24 is welded and fixed to the nozzle 35, the turbine vane main body 21 can be assembled without the nozzle 18 being thermally affected and destroying the single crystallinity.

The present invention is not limited to the above-mentioned embodiments, but can be applied to a low-pressure turbine blade, and various changes can be made without departing from the scope of the present invention. Of course.

[0036]

As described above, according to the method of assembling the turbine vane of the present invention, the excellent effect that the nozzle can be assembled without exerting a thermal influence on the nozzle can be obtained.

[Brief description of drawings]

FIG. 1 is a partially cutaway side view of an embodiment of the present invention.

FIG. 2 is a view taken along the line II-II of FIG.

FIG. 3 is a view taken along the line III-III in FIG.

FIG. 4 is a partially enlarged view of FIG.

FIG. 5 is a schematic side sectional view of a jet engine.

6 is a partially enlarged view of FIG.

FIG. 7 is a partially cutaway side view of a conventional example.

FIG. 8 is a view on arrow VIII-VIII in FIG. 7.

9 is a view on arrow IX-IX in FIG.

FIG. 10 is a partially enlarged view of FIG.

[Explanation of symbols]

 18 Nozzle 19 Outer band 20 Inner band 21 Turbine vane body 22 Cooling groove 23 Cooling plate 24 Insert 31 Opening 32 Nozzle hollow part 32 34 Inner edge 35 Lip part 36 Brazing part 37 Welding part

Claims (1)

[Claims] 1. A turbine vane main body comprising an outer band and an inner band for mounting and fixing at both ends of a hollow porous nozzle having a blade shape, which is formed around an opening of a hollow portion of the nozzle in the outer band. The porous cooling plate is brazed so as to cover the cooling groove portion, and the inner edge portion of the cooling plate along the opening is bent outward to form a lip portion, and the lip portion is provided to the nozzle hollow portion. A method for assembling a turbine vane, which comprises fixing the inserted hollow porous insert by welding.